The effect of multigrid acceleration implemented within an upwind-biased Euler method for hovering rotor flows is presented. The requirement to capture the vortical wake development over several turns means a long numerical integration time is required for hovering rotors, and the solution (wake) away from the blade is significant. Furthermore, the flow in the region near the blade root is effectively incompressible. Hence, the solution evolution and convergence is different to a fixed wing case where convergence depends primarily on propagating errors away from the surface as quickly as possible, and multigrid acceleration is shown to be less effective for hovering rotor flows. It is found that a simple V-cycle is the most effective, smoothing in the decreasing mesh density direction only, with a relaxed trilinear prolongation operator. Results are presented for multigrid computations with 2, 3, 4, and 5 mesh levels, and a CPU reduction of approximately 80% is demonstrated for five mesh levels.

Most service failures in fatigue start on a free surface, as in many cases the most severe stress conditions exist there and are of a bi-axial stress nature, such as a drive shaft under combined bending and twisting. This particular case of multi-axial stress fatigue, because of its importance in design, has seen considerable investigation, although several aspects of the stress situation complicate the analysis of the test results with regard to the development of a general criterion of fatigue failure under multi-axial stress. The principal stress axes rotate with variations in the ratio of bending to twisting and the shaft is subjected to a stress gradient. Only the extreme portions of the shaft in the plane of bending are under the maximum bending stress, whereas the complete surface of the shaft is subjected to the maximum torsion stress.

The problem of designing a three-dimensional contraction arose recently in the course of the projected construction of a new low-speed (200 ft. /sec), closed-return wind-tunnel at the Northampton College of Advanced Technology. A three-dimensional contraction was required of contraction ratio 7·4:1 to provide smooth low turbulence conversion to the working section speed within an axial length of 126 inches. The inlet and outlet (working-section) dimensions are shown in Fig. 1 and the sections are similar. It was decided to design the contraction on the basis of an equivalent axi-symmetric contraction whose inlet and outlet circular areas are equal to those of the corresponding irregular octagons and then adapt the derived contraction wall profile to the actual sectional contours. The resulting equivalent inlet and outlet radii are then 55·4 in. and 20·4 in. respectively. The problem was thus reduced to the design of an axi-symmetric contraction of given area ratio within a specified axial length, with favourable velocity characteristics for low turbulence flow in the working section. This latter specification implies the absence of adverse pressure gradients on the walls near the working section and a small percentual difference between axial and wall velocities at the inlet to the working section.

The finite element method is extended to the problem of finite plane strain of elastic solids. A highly elastic body subjected to two-dimensional deformations is represented by an assembly of triangular elements of finite dimension. The displacement fields within each element are approximated by linear functions of the local coordinates. Non-linear stiffness relations involving generalised node forces and displacements are derived from energy considerations. For demonstration purposes, the non-linear stiffness equations are applied to the problems of finite simple shear and generalised shear. For finite simple shear, it is shown that these relations are in exact agreement with finite elasticity theory. Convergence rates of finite element representations of these problems are briefly examined.

Coupling an unsteady flow-solver with a structural model offers the opportunity to simulate aeroelastic behaviour of wings and rotor blades. The moving and deforming surfaces resulting from unsteady simulations require deforming meshes during the simulation, and it is common to use simple interpolation of surface displacements and velocities onto the initial undisturbed mesh. However, aeroelastic simulations can result in large displacements and deformations of solid surfaces, and simple interpolation of perturbations results in poor grid quality and possible grid crossover. A new interpolation technique is presented which is still simple in that it is driven solely by surface motion, but represents rotational effects near the solid surface, to maintain grid quality there. Furthermore, the scheme is fully analytic, so is very cheap computationally and results in grid speeds also being available analytically. Results, in terms of unsteady grid motion and flow solution, show the scheme to be effective and efficient.

In this paper two numerical procedures are presented: the first algorithm allows for the determination of the geometric characteristics of the maximum efficiency propeller for a given operative condition and profile distribution along the blade; the output of this numerical procedure is the chord distribution and twist angle of the blade, together with its efficiency and its torque and thrust coefficients for the prescribed operative condition. The aerodynamic characteristics of the optimum propeller when operating in a condition different from the design one are obtained by a second algorithm that allows for the evaluation of the efficiency, the thrust and torque coefficients of a propeller of known geometry, when the blade pitch and operative condition are varied.

In the paper the formulation used for deriving the geometry of the optimum propeller and determining its performances when operating off-design is described in detail. The results obtained from the proposed propeller model have been validated by comparison with experimental data.

The computation of slowly moving Shockwaves is considered. A numerical error is introduced into the solution during an explicit computation, which results in the generation of spurious entropy. The generation of this spurious entropy and its effect on the pressure wave in Euler flows is examined, in one dimension by considering a moving shock in a constant area duct, and in two dimensions by considering an unsteady aerofoil flow. An efficient temporal discretisation is presented, whose advantages over the explicit scheme are two-fold. First, the numerical entropy generation due to shock motion is eliminated, and second, the CPU requirements for a two-dimensional unsteady computation are reduced by a factor of five from that required for the explicit scheme, for no loss of accuracy.

During the past two decades, with the advancement of avionics, control systems, design, product methods, and communication systems, UAVs have gained more interest from both civil and military customers. Hence, more researchers have dedicated their time to developing new applications for drones and improving their performance. In this study, a literature review on drone applications is conducted. It is explained that they can be more effective and beneficial when merged with preexisting systems, and thus making a system of systems. Further, we have classified UAV applications more comprehensively. This classification can be very useful for providing insights when designing new UAVs. Finally, previous works are discussed thoroughly and diagrams are developed to show the research focus in this field. These statistical diagrams can also be used as a gap analysis tool.

Longitudinal vortices have tremendous practical utility for flow control (control by vortices) and in some cases have exhibited tremendous potential for causing harm if uncontrolled (i.e. control of vortices is required). Vortex control has thus far been carried out via multitudinous approaches in an empirical fashion, aided by the essentially inviscid nature of much of longitudinal vortex behaviour. Further refinement and several applications of vortex flow control require knowledge regarding the detailed flow physics of longitudinal vortices such as transition, transitional flow regimes, turbulence structure and modelling, and interaction with shock waves, other vortices and surfaces. This paper summarises vortex control applications and extant techniques for the control of longitudinal vortices produced by bodies, leading edges, tips, and intersections.

This work presents a method for selecting the gain parameters of a C* control law for an aircraft’s longitudinal motion. The design incorporates various handling quality requirements involving modal, time- and frequency-domain criteria that were fixed by the aircraft manufacturer. After necessary model order-reductions, the design proceeds in essentially two-steps: stability augmentation system (SAS) loop design and control augmentation system (CAS) loop design. The approach partly relies on the use of guardian maps to characterise, in each case, the set of gain parameters for which desired handling quality requirements are satisfied. The approach is applied throughout the full flight envelope of a business jet aircraft and yields satisfactory results.

Climate change is increasing global-mean tropospheric temperatures, but the localised trends are uneven, including cooling the lower stratosphere and lifting the tropopause. The wind speeds are also being modified, both at the surface and aloft. A further effect, additional to wind and temperature alone, is of increasing fluctuations and severity of extreme weather. These are impacting air transport, and this will continue. The effects are known to include increased take-off distances where excess runway lengths exist and reduced payloads where they do not, increased en-route flight times, increased frequency and severity of encounters with clear air turbulence in some regions, changed patterns of wildlife — particularly bird — activity in some regions (potentially also for other anthropogenic reasons) are shifting locations of flight safety hazards, and increased burdens upon airport and associated infrastructure. There is increasing understanding and acknowledgment by companies and authorities of these effects and the importance of mitigating them, although this is not universal and there are as yet no universally understood best practices for air transport climate change mitigation.

This paper presents an overview of an advanced, conceptual wing-box weight estimation and sizing model for transport aircraft. The model is based on linear thin-walled beam theory, where the wing-box is modelled as a simple, swept tapered multi-element beam. It consists of three coupled modules, namely sizing, aeroelastic analysis, and weight prediction. The sizing module performs generic wing-box sizing using a multi-element strategy. Three design cases are considered for each wing-box element. The aeroelastic analysis module accounts for static aeroelastic requirements and estimates their impact on the wing-box sizing. The weight prediction module estimates the wing-box weight based on the sizing process, including static aeroelastic requirements. The breakdown of the models into modules increases its flexibility for future enhancements to cover complex wing geometries and advanced aerospace materials. The model has been validated using five different transport aircraft. It has shown to be sufficiently robust, yielding an error bandwidth of ±3%, an average error estimate of -0·2%, and a standard error estimate of 1·5%.

Results of a comparison of dynamic stall onset are presented, as assessed from low Mach number windtunnel data contained in the University of Glasgow's database and that from the well established Beddoes’ model. The model, which was originally developed to reconstruct higher Mach number dynamic stall characteristics, exhibited a lower stall onset incidence than that assessed from the windtunnel data. The differences and speculations on the physical reasoning underlying the two assessments are discussed. An addition to the Beddoes model is proposed which yields an improved reconstruction of the Glasgow windtunnel data.

Following on from the LUMINA and HERMES elements discussed in TN’s 11 and 12, this note analyses two further three-dimensional elements outlined in ref. 1. These new elements possess the desirable features of completeness and consequent invariance of the polynomials for the displacement fields. It has been shown in ref. 2 and stated in ref. 1 that third order complete polynomials yielding a second order or parabolic strain distribution fit into a tetrahedron element with 20 nodal points and 60 degrees of freedom, denoted within ASKA as TET 20 (Fig. 1).

Based on two-phase flow theory, an Eulerian method to simulate rime ice accretions on an aerofoil has been developed. The SIMPLE (semi-implicit method for pressure linked equations) algorithm on a collocated grid is employed to solve the governing equations for the airflow. In order to simulate droplets impinging on an aerofoil, a permeable wall is proposed to solve the governing equations for supercooled droplets. The collection efficiency and impingement limits are obtained from the droplets’ flowfield. The process of ice accretion is simulated using the assumption that ice accumulates layer-by-layer and the ice shape is predicted with the assumption that ice grows in the direction normal to the aerofoil surface. The rime ice accretions on a NACA0012 aerofoil at 0° and 4° angles-of-attack have been investigated and there is agreement between the simulated results and previously published experimental data. The change of the pressure coefficient along the iced aerofoil is also analysed.

The use of aerospikes at the nose of a hypersonic vehicle can substantially reduce its drag; however, the associated heat flux at the reattachment point is very high for turbulent flows. In this paper, the aerothermodynamics of a generic projectile model represented by a hemisphere- cylinder, fitted axisymmetrically with an aerospike at the nose is investigated numerically with a commercially available Navier Stokes Solver. The base model is a hemisphere-cylinder with a diameter of 40mm and an overall length of 70mm. An aerospike protrudes axisymmetrically at the nose of the base model and has a hemispherical cap from which another aerospike protrudes which again has a hemispherical cap, thus forming a double-disk aerospike. Compressible, axisymmetric Navier-Stokes Equations are solved for turbulent hypersonic flow of thermally perfect air with free stream conditions of Mach No 6·2 at standard sea level atmospheric conditions. The results for the model with double-disk aerospikes are compared with those of hemisphere-cylinder without aerospike. The results suggest that the use of double-disk aerospike can favourably reduce the turbulent reattachment heat flux along with suitable drag reduction of the main body.

As simulation has become an integral part of the overall life-cycle support of aircraft, the need for effective Virtual Engineering (VE) tools to support these activities has increased. FLIGHTLAB is a state-of-the-art, aircraft modelling and simulation software tool, that has been designed to address this need and is widely used in rotorcraft design, analysis, test and evaluation, and full-flight simulation applications. This VE tool supports the development and analysis of both fixed and rotary wing aircraft with an extensive library of modelling components which have been successfully used and validated in numerous, real-world applications. These components provide comprehensive modelling of aerodynamic, structural, control and propulsion disciplines. Analyses include performance, dynamic response, stability and control, airloads, and structural loads. Graphical User Interfaces and an interactive scripting language provide user-friendly operation. This paper describes the capabilities and validation activities that have been undertaken to support the development of the commercial VE toolset FLIGHTLAB over the last 20 years and discusses future rotorcraft challenges that could be addressed by enhancements to current generation VE tools.

Recently, transformable Unmanned Aerial Vehicles (UAVs) have become a subject of great interest in the field of flying systems, due to their maneuverability, agility and morphological capacities. They can be used for specific missions and in more congested spaces. Moreover, this novel class of UAVs is considered as a viable solution for providing flying robots with specific and versatile functionalities. In this paper, we propose (i) a new design of a transformable quadrotor with (ii) generic modeling and (iii) adaptive control strategy. The proposed UAV is able to change its flight configuration by rotating its four arms independently around a central body, thanks to its adaptive geometry. To simplify and lighten the prototype, a simple mechanism with a light mechanical structure is proposed. Since the Center of Gravity (CoG) of the UAV moves according to the desired morphology of the system, a variation of the inertia and the allocation matrix occurs instantly. These dynamics parameters play an important role in the system control and its stability, representing a key difference compared with the classic quadrotor. Thus, a new generic model is developed, taking into account all these variations together with aerodynamic effects. To validate this model and ensure the stability of the designed UAV, an adaptive backstepping control strategy based on the change in the flight configuration is applied. MATLAB simulations are provided to evaluate and illustrate the performance and efficiency of the proposed controller. Finally, some experimental tests are presented.

This paper reviews the research on Gurney flaps and related high lift trailing edge devices. It investigates aerofoil performances at Reynolds numbers Re ≅ 105 and below, both with the clean configuration and various Gurney flap sizes. The device height is optimised, and a semi-empirical formula linking flap height to free stream speed and aerofoil chord is proposed. The analysis shows that the optimal size of the device is always below the boundary-layer thickness at the trailing edge. Discussion of results includes analysis of hysteresis loops occurring in the L/D performances. These are mostly due to large changes in drag and small changes in lift, which occur when the aerofoil is restored to the reference angle of attack.

A procedure is developed for the analysis of plane stress problems when yielding occurs locally. The region is divided into triangular elements and the deformation is analysed on a step-by-step basis, using the matrix notation developed by Argyris. The simple expressions which are derived for the element properties are applicable with any stress-strain relations which are stable and time-independent. Simple numerical examples are given.